The Orbital Refueling System (ORS); force torque sensor; Plasma Motor/Generator (PMG) proof of function; voice controlled system; infrared intercommunications; superfluid helium on orbit transfer; laser docking sensor; and the Small Expendable Deployment System (SEDS) are summarized

Levin, G. M.

English

NASA Technical Reports Server

N 8 8 -  1081 1 
THE FLIGHT DEMONSTRATION 
PROGRAM AND SELECTION 
PROCESS 
G. M. Levin 
NASA Headquarters 
Washington, DC 
SPACE CONSTRUCTION CONFERENCE 
AUGUST 6-7, 1986 
5 
https://ntrs.nasa.gov/search.jsp?R=19880001489 2020-03-20T08:52:04+00:00Z
Introduction: 
In October 1983 the Office of Space Flight at NASA Headquarters 
established a new program office to: 
would support program objectives of the Office of Space Flight; give younger 
NASA engineers "hands on" experience with the development of flight hardware; 
lead to new business for the Space Transportation System (STS); and stimulate 
present customers of the STS towards more efficient use of the unique 
capabilities of the space shuttle. The Flight Demonstration Program was 
initiated by undertaking the development of six demonstrations. 
these demonstrations, the Orbital Refueling System was successfully flown on 
STS 41-C on August 29, 1984. 
and ACCESS, were performed successfully on STS 61-B on November 29 and 
Decemberl, 1985. These latter two demonstrations, are the subject of this 
Space Construction Conference. Therefore, no further discussions of 
EASE/ACCESS is included in this paper. 
undertake a series of developments which 
The first of 
The second and third Flight Demonstrations, EASE 
In the coming years these missions will be followed by demonstrations of 
a voice control system for the shuttle's closed circuit television system, an 
infrared intercommunications system, a plasma motor/generator proof of 
function demonstration, the implementation of a force feedback device on the 
shuttle's remote manipulator system, the transfer of super fluid helium on- 
orbit, a laser docking sensor and a small expendable deployment system. 
Flight Demonstrations are selected annually by the Flight Demonstrations 
Office in the Office of Space Flight. 
the Flight Demonstrations Office, proposals for NASA funded Flight 
Demonstrations are submitted to NASA Headquarters in July of each year. These 
proposals undergo a peer review with the final selection announced by the 
Associate Administrator for Space Flight in October. The Flight 
Demonstrations Program also participates with the Department of Defense (DoD) 
in the management and manifesting of DoD Flight Demonstrations. 
Responsive to a solication issued by 
6 
Typical areas of emphasis for Flight Demonstrations are: 
Space Construction/Large Space Structure 
Expendable Resupply 
System Improvement 
Satellite Servicing 
Tether Applications 
The proposed flight readiness dates for future OSF flight demonstrations 
are shown in Figure 1. 
FLIGHT READINESS SCHEDULE FOR OSF FLIGHT DEMONSTRATIONS 
APPROVED DEMOS: 
VOICE CONTROLLED CCTV 
INFRARED COMM 
PLASMA MOTOR/GENERATOR A 
DEXTEROUS END EFFECTOR 
SUPERFLUID HELIUM TRANSFER 
LASER RANGING SENSOR 
LOW G SLOSH CONTROL 
SMALL EXPENDABLE 
DEPLOYMENT SYSTEM 
PROPOSED DEMOS: 
LENS ANTENNA DEPLOYMENT 
STRUCTURES AND ASSEMBLY 
VERIFICATION EXPERIMENT 
SATELLITE REPAIR 
1987 
b 
bXl 
A 
1988 
F M A M J  J A S O N 0  
A 
r 
1989 1121314 
6 
SFC 
A 
A 
h 
‘;6 
1990 
1121j14 
T 
1991 xlx 
a? 
4RC 
7 
Approved F l i g h t  Demonstrations 
The O r b i t a l  Refue l ing  System ( O R s )  
The Orbi ta l  Refue l ing  System was the  first F l i g h t  Demonstrations Program 
sponsored by t h e  Office of  Space F l i g h t .  
develop and demonstrate the  equipment and procedures  f o r  a hydrazine f u e l  
t r a n s f e r  system; t o  develop and demonstrate the  t o o l s  needed t o  i n t e r f a c e  w i t h  
e x i s t i n g  sa te l l i t es  t o  accomplish f u e l  t r a n s f e r ;  and t o  develop and e v a l u a t e  
s p e c i f i c  procedures t o  r e f u e l  p r e s e n t  sa te l l i t es  of  t h e  Landsat type .  
The o b j e c t i v e s  of  t h e  ORS were t o  
The ORS c o n s i s t e d  of  f i v e  major c a t e g o r i e s :  the  s t r u c t u r a l  subsystem; 
t h e  f l u i d  subsystem; the  a v i o n i c s  subsystem; t h e  thermal c o n t r o l  subsystem; 
and the  EVA t o o l s .  The ORS tankage u n i t  ( w i t h  i ts f u e l  t anks ,  p ip ing ,  
va lv ing ,  heaters, and a v i o n i c s )  was mounted on a Mission P e c u l i a r  Experiment 
Support  S t r u c t u r e  (MPESS) i n  t h e  cargo  bay o f  the  o r b i t e r .  I n  t h e  
demonstration c o n f i g u r a t i o n  the  ORS was capable  of  m u l t i p l e  t r a n s f e r s  of  
approximately 200 pounds of hydraz ine  f u e l .  
d e l i v e r  approximately 500 pounds of  f u e l  t o  an  o r b i t i n g  sa te l l i t e  by a f u e l  
l i n e  connected w i t h  a hydraz ine  s e r v i c i n g  i n t e r f a c e  t o o l  set. The ORS was 
i n t e r f a c e d  t o  the  Orbiter g e n e r a l  purpose computer, and ana log  channels  were 
a v a i l a b l e  f o r  data monitoring. Cont ro l  of t h e  ORS f u e l  t r a n s f e r  was from the  
A f t  F l i g h t  Deck. During t h e  O R S  demonstrat ion fuel was t r a n s f e r r e d  through a 
f i x e d  p r o p e l l a n t  bypass as w e l l  as through a f u e l  l i n e  connec t ion  established 
by an  EVA crew member. 
The ORS could  be modified t o  
The EVA t o o l s  were the  unique hardware items r e q u i r e d  t o :  permi t  access 
t o  t h e  manual f i l l  va lve ;  provide  redundant seals f o r  crew s a f e t y ;  and c o n t r o l  
f l u i d  flow t o  the  sa te l l i t e .  The t o o l  set c o n s i s t e d  o f  seven items designed 
f o r  sa te l l i te  engagement, va lve  opening, and va lve  c l o s i n g .  .The f u e l  t r a n s f e r  
u n i t  and va lve ,  once engaged, became permanently attached t o  the  s a t e l l i t e ,  
t h u s  provid ing  a s t anda rd  i n t e r f a c e  f o r  r e f u e l i n g .  
Force Torque Sensor 
The g o a l s  of  the Force Torque Sensor (FTS) F l i g h t  Demonstration are t o  
demonstrate an  enhanced c a p a b i l i t y  of the  s h u t t l e  Remote Manipulator System 
(RMS) i n  a z e r o  lrgrl environment. 
g r a p h i c a l l y  d i s p l a y  the  f o r c e s  and t o r q u e s  which are developed a t  the  wrist 
j o i n t  o f  the RMS. 
The RMS/FTS system w i l l  measure and 
The RMS/FTS system des ign  w i l l  i n t e r g r a t e  new e lements  o r  components wi th  
e x i s t i n g  s h u t t l e  subsystems. The new e lements  w i l l  i nc lude  the  Force Torque 
Sensor ,  t h e  f o r c e  to rque  s e n s o r  e l e c t r o n i c s ,  t h e  s e n s o r  d i s p l a y  subsystem, and 
t h e  c a b l i n g  and c o n t r o l  pane l .  E x i s t i n g  s h u t t l e  subsystems t h a t  w i l l  be 
i n t e g r a t e d  wi th  t h e  RMS/FTS i n c l u d e  t h e  RMS, payload bay c a b l i n g ,  a c losed  
c i r c u i t  TV f l i g h t  moni tor ,  power and a i r  cond i t ion ing .  
The FTS housing w i l l  c o n t a i n  t h e  FTS e l e c t r o n i c s  and w i l l  be mounted as 
an  i n t e g r a l  p a r t  o f  t h e  RMS. The FTS w i l l  be i n s t a l l e d  between the  end of  t he  
RMS and the  base of t h e  s t anda rd  end e f f e c t o r .  The s e n s o r  d i s p l a y  system w i l l  
be l o c a t e d  i n  t h e  crew compartment and w i l l  comprise a microcomputer and 
8 
graphics generator. 
FTS and resolve the forces and vectors will be graphically displayed on a 
Shuttle TV flight monitor. 
The sensor display system will process signals from the 
Plasma Motor/Generator (PMG) Proof of Function 
The objective of this experiment is to provide order-of-magnitude 
verification of the key physical processes involved in utilizing the Earth's 
magnetic field, the velocity of the shuttle and a conductor either to generate 
a voltage or a mechanical force (electrical generator or motor). 
The results of this experiment would provide the first firm performance 
data to guide planning and development of hardware and experimental procedures 
for subsequent more extensive tests such as the Tethered Satellite System 
science experiments and engineering development systems. 
uncertainty in operation of a PMG system lies in the performance of the plasma 
current llbrushesll to complete the "armaturer1 to ionosphere current path. 
numerous variables involved in theoretical computation of the plasma physics 
processes result in considerable uncertainty in the results obtained. 
direct measurement, in orbit, of the induced VXB voltage, current, coupling, 
and ionospheric circuit impedance can provide adequate verification that the 
calculated effects actually occur. 
The primary 
The 
Only a 
Initial verification of these fundamental issues can be obtained prior to 
the availability of the Tethered Satellite System by flying the 200 meter/lO 
watt experiment. Adequate voltage and separation from spacecraft wake effects 
will be achieved by the 200 meter wire, without having to deal with 
stabilization and retrieval systems credible zeroth-order demonstration of the 
hollow cathode plasma coupling vital to PMG performance will be achieved. The 
deployed wire and far end electronics package will be jettisoned at the 
completion of the experiment. 
Voice Controlled CCTV Demonstration 
The Voice Controlled System (VCS) is a voice-actuated control system 
which can be used to operate spacecraft equipment when the operator's hands 
(and eyes) are preoccupied with the operation of other systems. 
increases the efficiency of crew operations. 
the VCS will be used by a crewman to operate the Orbiter television cameras 
while he is simultaneously using the Remote Manipulator System (RMS) t o  deploy 
or retrieve a payload. 
payload bay and the payload during the deployment operation, normally the 
crewman has to remove his hands from the RMS control handles and shift his 
visual attention to the television control panel to select cameras and adjust 
their view and focus. 
control words into his headset to effect these tasks. 
This 
To demonstrate this technique, 
Since the cameras provide differing views of the 
By using the VCS, he will only have to speak certain 
The VCS will provide control output for a command vocabulary of 100 
discrete words or phrases. It will respond only to the voice of the crewman 
who has previously recorded his voice patterns into a memory module which is 
inserted into the VCS at the time of the operation. 
All comands provided by the VCS will parallel those on the normal 
television control panel and can be overridden by the television control 
9 
switches. A fail safe feature will be provided to automatically remove the 
VCS from the control circuits when power is removed from the VCS. 
In addition to normal camera adjustments, the VCS will provide sequences 
of camera adjustments in response to a single command word input. 
sequences of adjustments will automatically position the camera to 
predetermined operating conditions without requiring the attention of the 
crew. 
These 
Infrared Intercommunications 
The Infrared Intercommunications (IR COMM) is a wireless means of 
communicating voice or data between two or more crewman or between crewmen and 
an associated wall-mounted unit which ties into the Shuttle or Space Station 
avionics. Information will be transferred by means of modulated infrared 
light radiated from tiny body-mounted diode light sources. This light will 
bounce off the walls and other objects in the cabin, essentially "floodingff 
the cabin volume and eliminating blind spots regardless of source location. 
The receiver will detect the light mounted anywhere within the cabin 
regardless of which direction the transmitter or receiver are pointed. 
concept has been demonstrated both in a laboratory environment and the shuttle 
mock-up. 
This 
A deficiency of the presently used wireless (radio) crew communications 
unit is its inability to meet communications security requirements without the 
use of body-worn encryptors. Infrared Communications would provide that 
security. The IR light is effectively attenuated by the Shuttle windows; the 
optical energy will not interact with the electrical signals of the Orbiter 
wiring. An optical system would be small, wireless, and light-weight, having 
less of an impact on the crew. Multiple users can be readily accommodated by 
additional channels. In addition to providing secure communications for the 
crew, IR could provide the transfer of high rate biological data from the crew 
to a wall unit, thereby replacing the existing tape recorders. This would 
allow real time data analysis while maintaining full crew mobility. 
The purpose of the flight experiment is to demonstrate the capability of 
the system to transfer voice and data without dropouts using the flooded 
volume technique. In addition, it will prove the adaptability of the system 
to a space vehicle environment which incorporates multiple users. 
Superfluid Helium On Orbit Transfer 
With this demonstration, we plan to transfer superfluid helium (at 
approximately 2OK) using 175 L dewars and under zero-g conditions. 
The method to be used involves a porous plug thermomechanical pump--a 
simple electrical device that, due to the unique properties of superfluid 
helium pumps the superfluid by simply heating one end of the propus plug. 
principle involved is the same a's that used by the, IRAS and COBE porous plug 
to contain superfluid helium within a dewar in zero-g. 
The 
Many current and proposed space flight projects include detectors, 
instruments (AXAF, COBE, LDR), and facilities (IRAS, SIRTF, GP-B) which are 
liquid helium cooled. For long lifetime instruments and facilities, the 
10 
l i m i t i n g  f a c t o r  is t h e  d e p l e t i o n  o f  the  s t o r e d  helium. For t h o s e  faci l i t ies  
which, wi th  re furb ishment ,  are requ i r ed  t o  last more than  1 o r  2 y e a r s  (SIFTF, 
AXAF, L D R ) ,  a method o f  r e p l e n i s h i n g  t h e  l i q u i d  helium is r e q u i r e d .  
We b e l i e v e  t h a t  t h i s  demonstration is a n a t u r a l  follow-on f o r  t h e  
hydraz ine  p r o p e l l a n t  r e f u e l i n g  demonst ra t ion  r e c e n t l y  completed. 
s u p e r f l u i d  helium t r a n s f e r  w i l l  demonstrate ano the r  i n - f l i g h t  s e r v i c i n g  
c a p a b i l i t y  f o r  t he  s h u t t l e .  
The 
We are c u r r e n t l y  s e t t i n g  up a l a b o r a t o r y  t r a n s f e r  experiment a t  the  
GSFC. T h i s  experiment w i l l  f u r t h e r  d e f i n e  t h e  porous p lug  parameters  r e l e v a n t  
t o  t h e  t r a n s f e r  o f  s u p e r f l u i d  helium i n  zero-g. 
Laser Docking Sensor 
T h i s  proposed f l i g h t  demonst ra t ion  is  designed t o  show the  a b i l i t y  o f  t h e  
Laser Docking Sensor (LDS) t o  provide  real  time data as an  a i d  i n  a manual 
docking procedure.  
dimensional target a t t i t u d e  s imula t ion  w i l l  be a v a i l a b l e  as a v i s u a l  d i s p l a y  
provided by t h e  f l i g h t  q u a l i f i e d  g r i d  computer. 
t h e i r  basis f o r  performing the  approach t o  t h e  target. 
provide  azimuth and e l e v a t i o n  a n g l e s  as well as range t o  r e t r o r e f l e c t o r s  
p laced  i n  a known c o n f i g u r a t i o n  on t h e  target v e h i c l e .  
Rendezvous and docking parameters  a long  wi th  a three- 
The crew may use t h i s  data as 
The LDS system w i l l  
By knowing these parameters ,  a p lane  can be de f ined  whose o r i e n t a t i o n  
r e p r e s e n t s  t he  a t t i t u d e  ( r o l l ,  p i t c h ,  yaw) of  t h e  target wi th  r e s p e c t  t o  t he  
s e n s o r / a c t i v e  v e h i c l e  axes .  T h i s  i n fo rma t ion ,  provided a t  a rea l  time rate 
(about  3-5 Hz), w i l l  enab le  an  optimum approach t r a j e c t o r y  t o  be c a l c u l a t e d  
and followed e i ther  manually, as a v i s u a l  a i d  t o  the  crew, o r  a u t o m a t i c a l l y ,  
by p rov id ing  the  informat ion  t o  the  e x i s t i n g  onboard nav iga t ion  computers. 
The LDS comprises a n  i l l u m i n a t i n g  laser source  t h a t  w i l l  e i ther  scan  over  
o r  f l o o d  the  f i e l d  of view. T h i s  w i l l  p rovide  t h e  high s i g n a l  t o  n o i s e  r a t i o  
r e q u i r e d  t o  main ta in  the  accuracy  o f  t h e  system. A r e t r o r e f l e c t o r  
c o n f i g u r a t i o n  w i l l  be a t t a c h e d  t o  any target v e h i c l e  i n  a known c o n f i g u r a t i o n  
o f  known dimensions t o  a l l o w  the  c a l c u l a t i o n s  of unambiguous a t t i t u d e .  
Ranging w i l l  be accomplished by modulating t h e  laser source  i n  e i ther  a pulsed  
f a sh ion  (pu l sed  o r  ch i rped  ranging)  o r  a CW f a s h i o n  ( t o n e  r ang ing) .  
technology w i l l  p rovide  range accuracy  t o  approximately one pe rcen t  o f  
range. 
from t h e  r e t r o r e f l e c t o r s .  
P r e s e n t  
A r e c e i v e r  w i th  accompanying o p t i c s  w i l l  detect t h e  modulated r e t u r n  
The o p e r a t i o n a l  s c e n a r i o  w i l l  c o n s i s t  o f  scanning the  f u l l  f i e l d  o f  view 
u n t i l  t h e  r e t r o r e f l e c t o r s  are encountered. Each r e t r o r e f l e c t o r ,  i n  t u r n ,  w i l l  
be l o c a t e d  i n  a n g l e  and range. The s e n s o r  w i l l  then  e n t e r  t h e  t r a c k  mode 
which c o n s i s t s  o f  performing small, fast d i t h e r  s c a n s  over  t he  l as t  known 
l o c a t i o n  of each r e f l e c t o r .  R e f l e c t o r  movements w i l l  be then  detected t o  
a l low the c a l c u l a t i o n  o f  a t t i t u d e ,  a n g l e ,  rate, range  and range  rate r e l a t i v e  
t o  t h e  a c t i v e  v e h i c l e .  
11 
Small Expendable Deployment System 
Th i s  proposed tether a p p l i c a t i o n s  F l i g h t  Demonstration w i l l  check-out and 
v a l i d a t e  the  Small Expendable Deployment System (SEDS) concept.  
demonstration the  te ther  w i l l  be c u t  and allowed t o  r e e n t e r  the Earth 's  
atmosphere after use.  
w i l l  be used t o  boos t  a 350 pound end mass about  70 km above the  s h u t t l e ' s  
o r b i t .  
e n t i r e  system w i l l  be mounted on a Mission P e c u l i a r  Experiment Support  
S t r u c t u r e  as a p a r t i a l  payload. 
I n  t h i s  
The proposed system w i l l  have a 10 km long tether which 
The tether w i l l  be housed i n  a Get Away S p e c i a l  (GAS) can and the  
SEDS w i l l  weigh approximately 450 pounds inc lud ing  t h e  end mass. The 
The tether w i l l  unwind from t h e  
diameter of t he  k e v l a r  tether w i l l  be 1.65 mm. 
wound on a spoo l  and placed i n  a GAS can. 
o u t s i d e  s u r f a c e  o f  the  spool .  
w i l l  be  a p p l i e d  t o  the  tether v i a  gas bags l o c a t e d  between t h e  spoo l  and the  
GAS can walls. 
t h e  proper  p r e s s u r e  i n  t h e  bags. 
us ing  an  o p t i c a l  t u r n  coun te r  and a c lock .  
frame b u i l t  above a GAS can w i l l  p r o t e c t  the  tether and s e r v e  as an e x i t  pa th  
f o r  t h e  tether.  
tether when t h e  experiment is completed. 
The concept c o n s i s t s  o f  k e v l a r  
For c o n t r o l  of  the  deployment speed ,  t e n s i o n  
A gas supply and meter ing  system w i l l  be needed t o  main ta in  
The speed of  deployment w i l l  be measured 
A converging-shaped s t r u c t u r a l  
A c u t t i n g  t o o l  l o c a t e d  a t  t he  t o p  of t he  frame w i l l  s e v e r  the  
12 


The flight demonstration program and selection process

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